Advanced geometry main rotor blade

ABSTRACT

A rotor blade for a convertiplane characterized by twist rate which is nonlinear but monotonic from blade root to blade tip and by camber nonlinear from blade root to provide a lift due to camber which decreases from blade root to an intermediate point along the length and increases from said intermediate point to blade tip.

United States'Patent [191 Edenborough et al.

[45'] Aug. 13, 1974 ADVANCED GEOMETRY MAIN ROTOR BLADE Inventors: Harry K. Edenborough, Dallas;

Kenneth G. Wernicke, Hurst;

George D. Carter, Fort Worth, all of Tex.

Assignee: Textron Inc., Providence, RI.

Filed: May 17, 1971 Appl. No.: 143,850

US. Cl. 416/223 Int. Cl. B64c 27/06 Field of Search 416/226, 223, 232, 233

References Cited UNITED STATES PATENTS 2/1952 Bensen 416/226 X 4/1958 Bath .1 416/226 UX 8/1960 Huber 416/226 3,065,933 11/1962 Williams 416/242 UX 3,167,129 l/1965 Shultz 416/226 3,173,490 3/1965 Stuart 416/223 3,501,248 3/1970 Brocker 416/89 3,552,881 l/l97l Rogers et a1 416/226 X 3,558,081 l/1971 Williams 416/242 X Primary Examiner-Everette A. Powell, Jr. Attorney, Agent, or Firm-Richards, Harris & Medlock [5 7] ABSTRACT A rotor blade for a convertiplane characterized by twist rate which is nonlinear but monotonic from blade root to blade tip and by camber nonlinear from blade root to provide a lift due to camber which decreases from blade root to an intermediate point along the length and increases from said intermediate point to blade tip.

7 Claims, 14 Drawing Figures Pmmznwm 3W 3,529,240

SHEET 2 DF 3 FIG 4 movsao-aaswvo INVENTORS;

HARRY K. EDENBOROUGH KENNETH G. WERNICKE GEORGE D. CARTER ATTORNEYS PMEMEQ am; i 3197 SHEET 3 or a L J i GgVWMiI l \L is L 1 g u. g 1 8 INVENTORS'. HARRY K. EDENBOROUGH KENNETH G. WERN/CKE GEORGE D. CARTER ATTORNEYS I ADVANCED GEOMETRY MAIN ROTOR BLADE BACKGROUND OF THE INVENTION Military and civilian air operations contemplate the availability of an aircraft which combines efficient high speed cruise with vertical take-off and landing (VTOL) capability. Military operations must be conducted in all kinds of terrain beyond the range from large air fields. Such aircraft would be of great advantage in military operations. Civil air transportation would be greatly enhanced by the availability of VTOL aircraft operating between city centers as much as 500 miles apart. They would serve to reduce congestion at regional airports by relieving such airports of short haul passengers. Operating between a city center and an airport, such aircraft would shorten the city to airport travel time for the impatient traveler.

In attempts to meet such needs, VTOL aircraft have heretofore been built and tested. The attempts to satisfy the requirements of a high payload vertical lift capability as well as high speed cruise efficiency has led to the development of the present invention of a main rotor blade of advanced geometry particularly suitable for a convertiplane.

RELATED APPLICATIONS The blade construction disclosed in the present application is described and claimed in (application Ser. No. 143,970, filed concurrently herewith.)

The blade geometry disclosed in U. S. application No. 143,970 is efficient in both the helicopter and the airplane mode. The blade is twisted in the sense that the Zero lift line changes in pitch with distance from the blade root. Further, an aerodynamic twist is achieved by changing the camber of the blade with distance from the root. A change of camber effectively changes the zero lift line of the blade.

In order readily to provide such a blade which may reliably be repeated, construction substantially different than heretofore known is preferably employed.

SUMMARY OF THE INVENTION In accordance with the present invention, there is provided a proprotor blade of unitary structure having both geometric twist and aerodynamic twist, the latter being achieved by varying camber. More specifically, geometric blade twist change is nonlinear and monotonic blade blade root to blade tip. Camber changes nonlinearly from root to tip to provide a design lift coefficient which decreases from blade root to an intermediate point along the blade and increases from the intermediate point to blade tip. Preferably, the blade twist is about 30 from root to tip while camber, in terms of design lift coefficient, decreases fromabout 0.7 at the root to 0.1 near mid length and then in creases to about 0.2 at the blade tip.

In a more specific aspect, the total effective twist is about 45 and results from a geometric twist of about 30 with effective twist from camber added thereto.

BRIEF DESCRIPTION OF THE DRAWINGS For a more complete understanding of the invention,

reference may now be had to the following description taken in conjunction with the accompanying drawings in which:

FIG. 1 is a view of a convertiplane to which the blade of the present construction is particularly adapted;

FIG. 2 illustrates the twist and camber of the blade;

FIG. 3 is a graph, details of which are represented by FIG. 2;

FIG. 4 is a sectional view of the blade taken along lines 4-4 of FIG. 6;

FIG. 5 is a sectional view taken along the lines 55 of FIG. 4;

FIG. 6 is a top view of a blade of FIG. 4;

FIG. 7 is a sectional view of the blade taken along lines 7--7 of FIG. 6;

FIG. 8 is a partial sectional view of the blade taken along lines 8-8 of FIG. 6;

FIG. 9 illustrates the manner :in which the blade is coupled to the main mast;

FIG. 10 is a view of the root end of the blade;

FIG. 11 illustrates the manner in which the spar is formed;

FIG. 12 illustrates fabrication and tooling of the blade with the preformed spar as an integral part thereof;

FIG. 13 is a graph illustrative of the hover performance of the blade; and

FIG. 14 is a graph illustrative of the propulsive efficiency of the blade in the airplane mode.

FIG. 1

Referring now to FIG. 1, a convertiplane 10 has been illustrated in which twin three-bladed proprotors, such as the rotor 11, mounted on a pylon 12 are provided for vertical take-off in the helicopter mode and for forward propulsion in an airplane mode of operation.

Preferably, the proprotor 11 is semirigid with the hub gimbal mounted to the pylon mast to provide blade flapping freedom with all of the functional requirements for varying blade pitch and. for retention thereof provided.

Fundamental understanding of the proprotor/pylon phenomena has heretofore been gained through construction operation of such craft as the XV-3 convertiplane. Such operations led to further studies, the emergence, and verification of the present invention.

FIGS. 2 and 3 The geometry of a proprotor of the present invention is illustrated by FIG. 2 which portrays the twist of the blade and variation in camber. FIG. 2 corresponds with the criteria illustrated in the graphs in FIG. 3. The blade having the configuration illustrated in FIGS. 2 and 3 incorporates the desirable characteristics of a helicopter blade and an airplane propeller. A blade having characteristics illustrated in FIGS. 2 and 3 functions efficiently in its dual role as a helicopter rotor and an airplane propeller. By the present invention the blade leads to an efficient lightweight proprotor. A blade having the combination of geometrical twist and camber illustrated in FIGS. 2 and 3 meets the aerodynamic requirements for both helicopter and airplane flight. In such a blade the blade spar structure may have a uniform twist rate and an integral grip. This eliminates the need for an aerodynamic cuff at the root of the blade and at the same time saves weight and minimizes performance losses.

Twist as used herein refers to the spiral configuration of the blade. Camber is a direct measure of the lack of symmetry of the blade cross section relative to the chord line. If the blade cross section is symmetrical to the chord line, it has zero camber. If the distance from a given point on the chord line to skin surface in one direction is greater than in the other, then camber is present. Camber may be said to introduce an effective twist by shifting the airfoil zero lift line which may add to or subtract from the geometrical twist in control of the total effective twist of a blade and thus tailor the blade to a predetermined performance criteria.

For example, in the embodiment illustrated in FIG. 1, the proprotor system is about 25 feet in diameter. In FIG. 2, a preferred blade geometry is shown. The perimeter and chord line is shown for each of the 31 slices through the blade. The slices are at 5 inch intervals in a blade 150 inches in length.

The blade chord at the root is represented by the line 20. The second blade chord is represented by the line 21. The thirtieth blade chord is represented by the line 22 and the thirty-first blade chord is represented by the line 23. The intermediate blade chords for the other points along the blade are illustrated. The angular variation in the blade chord line indicates that the blade is twisted by about 45from the root to the tip. From FIG. 3, line 24 shows that the rate of twist with distance is not constant.

Curve 20a corresponds with the outer surface of the blade and thus represents blade configuration at the root. The curve 21a represents the cross-sectional outline of the blade as would be viewed if the blade were sliced 5 inches from the root. In a similar manner, the curve 22a represents the blade configuration 5 inches from the tip. The curve 230 represents blade configuration at the tip.

As above noted, the transition from root to tip is illustrated by the curves of FIG. 3. The geometric twist is illustrated by the curve 24 and the camber is represented by curve 25. Both curves 24 and 25 are nonlinear with distancefrom the root. Twist angle is ploted, in degrees, as a function of percentage of the blade length with the zero point taken at the mast center. Geometric twist changes linearly from about 30 to about at a point about 40 percent of the proprotor radius. The twist then changes linearly at a different rate to about 1 at the tip. The change of twist is therefore monotonic, being linear inboard and outboard of the inflection point at about 40 percent of the distance from mast to blade tip.

As to camber, variations in design lift coefficient (C are plotted as a function of blade length. The camber,'in terms of design lift coefficient, decreases from about 0.7 at the root to about 0.1 at the 40 percent length point and then increases to about 0.2 at the tip. Curve 25 is not monotonic, but rather exhibits a change of sign of slope about midway of blade length. The change in camber contributes an effective twist. The sum of the geometric twist and the twist due to camber is represented by curve 26, wherein the total effective twist varies monotonically but nonlinearly from about 40 to zero with the major change in the first 40 percent of blade length. This permits the blade spar structure to have a uniform twist rate while optimizing performance characteristics in both the helicopter and airplane mode. FIG. 3 also illustrates length to thickness variations in the blade. Thickness of 27 percent, 18 percent, 12 percent and 8 percent of chord length are found at about 20 percent, 52 percent, percent and percent points of the proprotor radius.

FIGS. 4-10 Referring to FIG. 4, a blade having the foregoing characteristics may be formed of two basic elements, a spar 30 and an after body section 31.

Spar 30 preferably comprises an abrasion strip in the form of a curved or U-shaped channel 33. Adhesively secured inside strip 33 are two spar reinforcing channels 34 and 35. The channels overlap in zone 36 to form, with the abrasion strip 33, a front cavity containing a nose block 37. The other flanges of the channels 34 and 35, namely ends 38 and 39, are short and are inwardly turned. A spar web panel 40 is positioned between channels 34 and 35 and is adhesively secured in abutment against the ends 38 and 39.

Spar web panel 40 is shown in the enlarged sectional view of FIG. 5. Panel 40 comprises a front skin 41 and a rear skin 42 with a honeycomb spacer unit 43 wherein the honeycomb cells extend parallel to the blade chord 30a. As shown in FIG. 4, the spar 30 further includes an internal doubler strip 44 on the upper inside surface of channel 35. Further, external doubler strips 46a and 46b are stacked and secured on the upper outer surface of abrasion strip 33. Doubler strips 470 and 47b are stacked and secured on the lower outer surface of the strip 33.

After body unit 31 includes an upper blade skin 50 and a lower blade skin 51. A honeycomb unit 52 is adhesively secured at the upper end thereof to the inner surface of the upper skin 50. A honeycomb unit 53 is adhesively secured to the inner surface of the lower blade skin 51. Cells of the honeycomb units 52 and 53 are oriented substantially perpendicular to the blade chord 30a. A trailing edge construction 54 joins the trailing edges of the skins 50 and 51. A strip 55 of adhesive secures the rear face of the spar web panel 40 to the front of the honeycomb units 52 and 53.

The honeycomb units 52 and 53 are intercised at the plane of the blade chord. The intercised ends of the honeycomb cells are then adhesively secured to each other.

The structure is a unitary all bonded blade with minimum weight. The abrasion strip 33, channels 34 and 35, skins 41 and 42, doublers 44, 45, 46 and 47, and skins 50 and 51 preferably are all of heat treated (17-7 PH) stainless steel. In the above embodiment, the spar web panel skins 41 and 42 and the blade skins 50 and 51 were of 0.008 inch thickness. The abrasion strip 33 was of 0.040 inch thickness and the channels 34 and 35 were of 0.032 inch thickness.

FIG. 6

In FIG. 6, the blade has been shown in plan view. An extension 60 of the spar 30 receives a root fitting. Such fitting secured in the end of the spar 30 becomes a unitary part thereof as will be further explained. In the embodiment illustrated, the doubler panels 46a and 46b both extend substantially symmetrical with respect to the line 30b which is the line about which the blade pivots to change pitch. Doubler 46a also has a wing portion dfic extending away from the root to the trailing edge of the blade. Similarly, the doubler 46b has a wing portion 16d extending to the trailing edge of the wing nearer the root.

A fiberglass fairing 54a is bonded onto the trailing edge of the blade. Te fairing 54a is shown in section in FIG 4.

FIGS. 7 and 8 FIG. 7 shows the after body skins 51 and 52 sealed together at the trailing edge. Further, a metallic nose block 37a is provided. At the section line 7-7, FIG. 6, the blade is much thinner than the section shown in FIG. 4. The section shown in FIG. 4 was taken 37.5 inches from the root end of the blade, whereas the section of FIG. 7 was taken 1 12.5 inches from the root of the blade.

In FIG. 5 the spar only is illustrated with both a nose block 33a and a spar filler block 33b.

FIGS. 9 and 10 The manner in which attachment is made to the rotor mast is illustrated in FIGS. 9 and 10. It will be noted that a root fitting 70 is bonded into the end of the blade with ends of the abrasion strip 33 extended to enshroud the fitting 70. Strip 33 terminates at the end of the fitting 70. Te fitting 70 is adhesively secured in spar 30. Fitting 70 has a tapered inner recess into which a spindle 71 extends. The blade pitch motion is accommodated by suitable needle bearings between spindle 71 and fitting 70. Blade retention is by means of a conventional wire strap unit having a strap fitting 72 with the wires extending around a spool 73 through which a strap pin 74 extends. A yoke assembly 75 having three spindles including spindle 71 is mounted for rotation with the main mast 77 by way of a universal joint assembly 78.

Preferably the yoke assembly is made out of titanium and is of stiffness such thqt all in-plane and out-ofplane blade frequencies are above rotor speed except the first (rigid body) blade flapping frequency which is below one per revolution as a result of the positive pitch-flap coupling utilized to stabilize blade motion in the airplane mode of flight. Te latter coupling and the manner in which it is utilized may be as described and claimed in U. S. Pat. No. 3,494,706..

FIG. 9 is a partial end view showing the abrasion strip 33 with fitting 70 mounted therein. Fitting 70 has a central opening 80 into which the spool 73, FIG. 10, is to be inserted with the strap pin 74 extending through the blade along the strap pin center line 81.

FIGS. 11 and 12 In order to fabricate the system, it will be understood as above indicated that the parts are secured by bonding with adhesive of the heat setting type.

The first fabrication step may be to form the spar 30 from the assembly of the abrasion strip 33, spar channels 34 and 35, inner doublers 44 and 45, and the spar web panel 40. The foregoing members are assembled and placed into a suitable fixture as shown in FIG. 11. Internal pressure bags 90 and 91 are then placed back of the panel 40 and inside of the spar 30, respectively. The pressure bags are then inflated to force the members into contact with the preformed surfaces of the bonding tool halves 92 and 93. The spar elements within the bonding tool are then placed in an autoclave where they are heated to cure the adhesive previously applied to the various confronting surfaces.

In the next step of fabrication, the honeycomb slices suitably shaped to conform with the approximate desired outer contours of the blade are adhesively secured to the inner surfaces of the upper and lower blade skins.

With the honeycomb halves 52 and 53 secured to skins 50 and 51, respectively, the spar 30 and the two skin units along with the outer doublers are placed in a bonding fixture, such as illustrated in FIG. 12, and placed in an autoclave.

The honeycomb sections 52 and 53 are made of height such that they interfere at the chord plane. The elements 95 and 96 of the bonding fixture are forced together as the fixture is closed. The cells of the honeycomb units 52 and 53 will be misaligned as they are brought together. As pressure is applied to the members 95 and 96, the ends of the honeycomb become intercised at the blade chord plane. That is, the confronting ends of the honeycombs will cross one another at angles. When they are forced together, one part will slice its way into or through an opposite part. The reaction outwardly of the honeycomb halves onto the forms 95 and 96 by the intercising action forces the upper skin 50 and the lower skin 51 into contact with the inner surfaces of the tools 95 and 96. By this means exact conformity of the blade contour to the form surfaces is assured. A bag 98 located internally of spar 30 forces the spar into engagement with the tool surfaces so that the external doublers will be forced intimately against the surface of spar 30 and over the leading edges of the skins 50 and 51. It will be noted that the skins 50 and 51 extend over the trailing edge of the FIGS. 13 and 114 Results obtained through the use of the present invention are illustrated in FIGS. 13 and 14.

FIG. 13 has been included herein to illustrate the correlation of the hover performance with the theory involved. In FIG. 13 the thrust coefficient, C has been plotted as abscissa with ordinates in terms of power coefficient, C p. The theoretical relationship between power coefficient and thrust coefiicient is represented by a solid line 121. The actual measured values of thrust coefficient are represented by the points denoted by squares and the points denoted by triangles. In test, the mast of the rotor was tilted at an angle of from horizontal. The data represented by the squares in FIG. 13 was with a tip speed of 600 feet per second. The data represented by the triangles were taken at a tip speed of 740 feet per second. It will be noted that at the point where the blade would start to stall, i.e., at point 122 on curve 121, the measured thrust was about 15 percent in excess of the predicted maximum thrust. This point is represented by point 123.

In FIG. M, the graph illustrates the correlation of propulsive efficiency of the aircraft in the airplane mode with theory. A calculated relationship is represented by the solid line 130. The actual test results at 600 feet per second tip speed are represented by the triangles. The test results at 400 feet per second tip speed are represented by squares. The test results at 500 feet per second tip speed are represented by circles, and tests at 740 feet per second tip speed are represented by diamonds.

Propulsive efficiencies in excess of 90 percent will be noted for a considerable number of the test points.

Since the thrust required in the airplane mode is less than the thrust required in the helicopter mode, the design involves a cruise tip speed of 600 feet per second. In contrast, the total lifting power required in the helicopter mode suggests operation at the higher speed of 740 feet per second. FIG. 14 confirms correlation between the test data and theory with the theory having been generally conservative. The graph of FIG. 14 shows that the propulsive efficiency of 78 percent may be realized in level flight at this speed. The prop rotor would have an efficiency of about 92 percent in full power climb.

By the present invention, the blade of complex shape, having a nonlinearly varied twist and camber is provided. In this blade geometrical twist and camber both change nonlinearly with blade length. The geometrical twist changes monotonically and camber changes nonmonotonically with an inflection point at about the same point as the major change in twist rate.

While other construction methods and different components may be used to make blades embodying the present invention, the structure preferable will be as illustrated in FIGS. 4-10.

Having described the invention in connection with certain specific embodiments thereof, it is to be understood that further modifications may now suggest themselves to those skilled in the art and it is intended to cover such modifications as fall within the scope of the appended claims.

What is claimed is:

l. A blade for a convertiplane proprotor comprising a unitary structure having twist and camber both of which vary nonlinearly along the length of the blade from blade root to blade tip with the geometrical twist decreasing linearly from root to near mid length and decreasing linearly but less severely from near mid length to tip and with the camber decreasing from blade root to near mid length and increasing from mid length to tip.

2. In a convertiplane, a proprotor blade having camber and twist both varying nonlinearly relative to blade length with camber decreasing from blade root to an intermediate point along the length and increasing linearly from said intermediate point to blade tip to permit the blade spar structure to have uniform twist rate while optimizing performance characteristics in both helicopter and airplane mode.

3. The combination set forth in claim 1 wherein said twist is about 30.

4. The combination set forth in claim 1 wherein the blade thickness to chord length ratio varies from about 0.27 near the blade root to about 0.08 at the blade tip.

5. The combination set forth in claim 1 wherein camber varies in terms of design lift coefficient from about 0.7 near the blade root to about 0.1 at about the 40 percent length point to about 0.2 at the blade tip.

6. A blade for a convertiplane proprotor comprising a unitary structure having twist therein varying nonlinearly from blade root to blade tip and having camber varying nonlinearly from blade root to blade tip to provide a lift coefficient decreasing from blade root to an intermediate point along the length and then increasing from said intermediate point to blade tip.

7. A blade for a convertiplane proprotor comprising a unitary structure having twist therein varying nonlinearly from blade root to blade tip and having camber varying nonlinearly from blade root to blade tip to provide a camber lift coefficient, decreasing from blade root to an intermediate point along the length and increasing from said intermediate point to blade tip with blade thickness to chord length ratio of from 0.27 near said root to 0.08 at said tip.

Poiwso UNITED STATES PATENT OFFICE CERTIB GATE OF CORRECTION Patent: No. 38292u0 Dated August 3, 9-7 1 lnvencofl Harry K Ede'nborough; K nneth 'G. Wernieke and a George D. Carter It is certified that error appears in the aboi7e-identified patent and that said Letters Patent are hereby corrected as shown below:

t" a '1 Col. 1, line 56, "blade blade'f should be -ybetween blade--.

Col. line 29 "channel 35" should be -channel 3 1 and a doubler strip M5 on the lower inside surface of the channel 35-.

C01. 5, line 6, "Te'? should be --T line .27, "Te" should be ----The--; line 39, "thqt" should be --thatline L 1, "Te" should be --The--.

Signed andsealed this 3rd day of December 1974,

(SEAL) Attest:

McCOY M. sms oN hR. CQJMARSHALL 1mm Attesting Officer Commissioner of Patents 

1. A blade for a convertiplane proprotor comprising a unitary structure having twist and camber both of which vary nonlinearly along the length of the blade from blade root to blade tip with the geometrical twist decreasing linearly from root to near mid length and decreasing linearly but less severely from near mid length to tip and with the camber decreasing from blade root to near mid length and increasing from mid length to tip.
 2. In a convertiplane, a proprotor blade having camber and twist both varying nonlinearly relative to blade length with camber decreasing from blade root to an intermediate point along the length and increasing linearly from said intermediate point to blade tip to permit the blade spar structure to have uniform twist rate while optimizing performance characteristics in both helicopter and airplane mode.
 3. The cOmbination set forth in claim 1 wherein said twist is about 30*.
 4. The combination set forth in claim 1 wherein the blade thickness to chord length ratio varies from about 0.27 near the blade root to about 0.08 at the blade tip.
 5. The combination set forth in claim 1 wherein camber varies in terms of design lift coefficient from about 0.7 near the blade root to about 0.1 at about the 40 percent length point to about 0.2 at the blade tip.
 6. A blade for a convertiplane proprotor comprising a unitary structure having twist therein varying nonlinearly from blade root to blade tip and having camber varying nonlinearly from blade root to blade tip to provide a lift coefficient decreasing from blade root to an intermediate point along the length and then increasing from said intermediate point to blade tip.
 7. A blade for a convertiplane proprotor comprising a unitary structure having twist therein varying nonlinearly from blade root to blade tip and having camber varying nonlinearly from blade root to blade tip to provide a camber lift coefficient, decreasing from blade root to an intermediate point along the length and increasing from said intermediate point to blade tip with blade thickness to chord length ratio of from 0.27 near said root to 0.08 at said tip. 